In gas turbine engines, air is compressed at an initial stage, then is heated in combustors, and the hot gas so produced drives a turbine that does work, including rotating the air compressor.
A number of existing gas turbine engine designs utilize some of the air from the air compressor to cool specific components that are in need of cooling. In some designs air is passed along a surface to provide convective cooling, and the air then continues to an intake of a combustor, and into the combustor where the oxygen of the air is utilized in the combustion reaction with fuel. This approach generally is referred to as “closed cooling.” In other designs, generally referred to as “open cooling,” air for cooling is passed into the flow of hot gases downstream of the combustion intake. In the latter cases a percentage of oxygen in such air for cooling may not be utilized in combustion, and this represents a potential inefficiency in that a percentage of the work to rotate the compressor does not supply air to the combustor intake for combustion purposes. The ultimate determination of whether it is more cost-effective to provide open cooling depends on balancing a number of factors, including expected component life cycle, and the costs of alternative cooling.
Combustor liners help define a passage for combusting hot gases immediately downstream of swirler assemblies in a gas turbine engine combustor. The surfaces of combustor liners are subject to direct exposure to the combustion flames in a combustor, and are among the components that need cooling in various gas turbine engine designs. An effusion type of open cooling has been utilized to cool combustor liners. This generally is depicted in FIG. 1A, which provides a cross-sectional of a prior art combustor 100. A predominant air flow (shown by thick arrows) passes along the outside of combustor 100 and into an intake 102 of the combustor 100. Centrally disposed in the combustor 100 is a pilot swirler assembly 104, and disposed circumferentially about the pilot swirler assembly 104 are a plurality of main swirler assemblies 106. Combustion generally takes place somewhat downstream of the pilot swirler assembly 104, designated in FIG. 1A as combustion zone 108. A transversely disposed base plate 110 receives downstream ends of the main swirler assemblies 106, and provides a physical barrier to flames that may otherwise travel upstream. An outlet 111 at the downstream end passes combusting and combusted gases to a transition (not shown, see FIG. 3).
Surrounding the combustion zone 108 is an annular effusion liner 112, and further outboard is a cylindrical frame 114. Welded to the frame 114 at its downstream end is an assembly of spring clips 116, which contacts a transition ring 120 of a transition (not shown in FIG. 1A). A plurality of holes (not shown) in the frame 114 allows passage of a quantity of air (shown by narrow arrows) that may pass through spaced apart effusion holes (not shown in FIG. 1A) in the effusion liner 112. FIG. 1B provides an enlarged view of the encircled section of FIG. 1A, in which spaced apart effusion holes 122 are depicted. The passage of air through the effusion holes 122 provides for a cooling of the effusion liner 112.
Referring to FIG. 1B, passage of air also is designed to occur along a radial gap 125 between the respective downstream ends 113 and 115 of the effusion liner 112 and the frame 114. The gap 125 is required to accommodate axial and radial differential expansion between the effusion liner 112 and the frame 114, and air flowing through the gap 125 also provides a cooling effect for the end of the effusion liner 112 and the frame 114. In certain embodiments a plurality of spaced apart protrusions 116 disposed at or near the end 113 of the effusion liner 112 establish the radial height of the gap 125.
Based on observation and analysis of present systems, such as that described in FIGS. 1A and 1B, and potential problems in some units of such systems, there is a need for an improved combustor liner that overcomes such problems.